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How significant is the lift induced drag contribution to total drag of aircraft at cruise speed?
How significant is the lift induced drag contribution to total drag of aircraft at cruise speed? From what I've learned so far it seems to be around 10 to 15 percent of total drag. I suspect the ratio doesn't vary too much between aircraft provided they're within the same speed range. Therefore, let's focus on airliners and other high subsonic jets.

There is a lot of discussion about optimum speeds in this question. The optimum cruise speed is a bit higher that the lowest drag speed—where induced drag is exactly half of drag—but it does not look like enough to only leave 15% induced drag.

These sources show 30-35% induced drag for transport aircraft at speed.
Here is a breakdown of 'typical' transport cruise drag from an analysis done for the erica tiltrotor. It shows about 35% induced drag at high speed cruise.

The Piano airframe analysis tool is generally regarded as representative of subsonic conventionally configured aircraft, and it shows 30% induced drag for the a320-200.


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Is wing sweep needed on supersonic aircraft?
Area rule states that bodies with the same crossectional area distribution have the same drag. Does that mean I could use a straight wing on mach 1.2 aircraft with performance as high as sweep winged plane as long as they have the same area distribution?

Sweep is not strictly required, but helpful to increase supersonic L/D. Just witness the length to which aircraft manufacturers went to sweep wings for Mach 2+ flight. However, sweep comes with its own set of problems.

A little more on the F104 vs sweept wings. In order to optimize a design for a given mach number greater than 1.0, the wing needs to be entirely enveloped within the mach pressure wave cone. Hence, large area with sweep or small area short and straight. Conversely, a long straight wing would be partially inside the mach cone and partially outside (not optimum).

Wing sweep is not required on supersonic aircraft. The F104 was a mach 2 interceptor with a straight, but tapered wing. It wasn't a well respected aircraft.

The Lockheed tests, however, determined that the most efficient shape for high-speed supersonic flight was a very small, straight, mid-mounted, trapezoidal wing.

Wow, I was right! Thank you!

It seems that swept wing is preferred on aircraft that are expected to fly a lot at high subsonic speeds, but aircraft optimized for pure supersonic can have straight wings

The ideal supersonic airfoil is a flat sheet of paper.  The 104's airfoil is "biconvex", a flat sheet with a bit of a curve on both surfaces.  It depended on a blown flap - bleed vented along the front of the flap to reduce flow separation when they are down, to get the approach speed down to a mere 160-170 knots or so, only about 30-40 kts faster than an airliner.  Not bad considering how small the wings are and how close they are to being flat sheets of plywood.

The 104 is, incredibly when you think about it, the basis for the U2 spyplane.

You will generally want to have a Swept Wing for supersonic aircraft as the swept wing delays the formation of shock waves.  The type of drag that a swept wing prevents is Wave Drag and would be a major source of drag at supersonic speeds.  I believe that the "Area rule" that you are quoting refers only to parasitic drag.  You may want to read this Wikipedia article on swept wings to learn more about their benefit.  The main downside of using swept wings is wing tip stall but that can usually be overcome with a Washout design.

I thought area rule referred to form drag, the energy used for displacement of the air by the body.  Parasitic is more like drag from bumps and projections etc no?

Wave drag and skin friction are both a type of parasitic drag

Form drag also is parasitic

Sweep or not, you will have wave drag from transsonic speeds upwards.

Another example of a supersonic aircraft that did not have swept wings is the Bell X1, the first aircraft to achieve supersonic speeds (and survive the experience). Given that the aircraft was pushing the unknown, the fuselage was shaped like a .50 caliber bullet, an object known to go supersonic smoothly... trial and error was a factor in pre computer aircraft design.
Note, though, that like the F104, the X1's wings are fairly short and set back on the fuselage, so that they are inside the shock wave cone coming off of the nose at supersonic speeds. As the X1 could barely go beyond Mach 1, its wings weren't as narrow as the F104, whose top speed was Mach 2.

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Could a kerosene fueled jet engine run on methanol without significant modifications?
Could a kerosene fueled jet engine run on methanol without sacrificing performance other than specific impulse? Would it require changes to the fuel injection/turbopump system? Would the combustor require a redesign? Consider both turbo engines and ramjets
(this question was previously posted on space SE due to a mixup, I've deleted it now)

Methanol is pretty hard on elastomers. I would be worried that any hoses or gaskets in contact with the methanol would fail. Plus the energy per mass is much lower with methanol because of those heavy oxygens already in it. You would probably have to recalibrate the fuel feed laws.

Related

You couldn't have any aluminum components in the fuel system because methanol dissolves aluminum (F1 race cars remove and purge their fuel systems right after racing).  Ethanol would be much better in that regard.
It would take changes to the fuel controller to change the air/fuel ratios used with less energy density.
Other than that, I don't think it would be a big problem.  I've been told that the Honeywell/Lycoming turboshaft in the M1 Abrams tank will run just about on any liquid hydrocarbon.. gasoline, kerosene, diesel, booze, rubbing alcohol...

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Why is Mode C helpful for ASDE-X systems?
Airports that operate an ASDE-X system require aircraft to operate transponders in mode C while moving on the surface.
While I understand the general purpose of operating the transponder to correlate with the other information ingested by the ASDE-X system, why is altitude (encoded in mode C) important for tracking surface movement?
Wouldn't mode A suffice?

Problem - many airliners do not transmit Mode A or Mode C when they're on the ground, it's inhibited. Does the ASDE work off Mode S ?

Without knowing the altitude, it is very difficult (or impossible) to accurately calculate the position of a transponder by using multilateration (which is what ASDE-X uses). In addition, without knowing the altitude, ASDE-X would be unable to determine whether an aircraft is on the airport's surface, or passing overhead.


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PT-6 Starting problems - PT6A - 67B, installed on a PC12
we are having trouble starting our PC12. After introducing fuel at stabilized 13% ng, the engine spools up very slowly - it takes forever to get through 20, 30, 40%ng. We've had several hung starts, and aborted starts.
Battery discharge test is fine - 106% on the first test, and it is a new battery.
Any suggestions for next step trouble shooting? Our mechanic doesn't seem to have a clue.

Our mechanic doesn't seem to have a clue. - Get a better mechanic.

A key piece of missing information:  Are you getting light-off?!

Either the starter generator isn't making the torque it should be making, or there is a fuel flow issue.  If I was betting 10 bucks on the result, I'd go with weak starter/generator and swap it out if another was available, or do diagnostic testing on the unit like insulation resistance checks and other electrical tests on the motor.
Regardless, if your mechanic is not sure how to proceed he/she should be on the phone to Pratt & Whitney Canada tech support or going through the PWC online portal to access the Fault Isolation Manual for that engine and just follow the steps.
If you're a pilot standing by watching the action, you'll really piss off the mechanic by relaying him troubleshooting instructions from some anonymous nobody from ASE, so I would only intervene if the mech was hiding in the bathroom and refusing to come out.


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How to reproduce NASA/Langley airfoils for general aviation with long chord lengths?
I have scaled a NASA/Langley LS(1)-0413 airfoil to a 90 inch chord length, and now the trailing edge is 0.5in. Any longer chord length would have a larger trailing edge thickness.

(90in*(0.0071-0.0016)=.495in)

This doesn't seem right, I have not witnessed a non-supercritical airfoil with a trailing edge larger than a .25in.
First off, do long chord length wings using the low and medium-speed airfoils have thick blunt trailing edges?
Secondly, if medium speed wings don't have blunt trailing edges how does industry taper the trailing edge? Is it assumed they will be rounded off, or sanded to a point? Should one remove from the bottom concave surface or remove from the top surface? Both will alter the exit angle of the airstream.
Thirdly, is there an ideal length chord for the NASA/Langley low and medium-speed airfoils, e.g. do they really only apply to a 36in chord lenght? Does scale matter?

Average chord is 65in, intended cruise is 250kts.

You are right, such a thick trailing edge is not right. It will add pressure drag unnecessarily – quick fix: Just taper the last 10 or 20 percent of the airfoil such that a thickness remains which is sufficient for structural reasons. You don't want a trailing edge that is as sharp as a knife edge. It has no stiffness left, so it will warp and will easily be damaged.
The more elaborate way: Do as if there would be a hinge at the nose and you close the two sides, just like scissors. That will leave the contour unchanged. The "hinge" location is ideally the center of the nose radius.
Tapering the last 10 or 20% will steepen the pressure rise there and will let the airfoil stall a tiny fraction of a degree earlier than with the unchanged surface contour.


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By what angle does the rotor disc change with max. cyclic input?
If maximum longitudinal/lateral cyclic blade pitch input is given, by how many degrees does the thrust vector (or rotor disc) change? Is the magnitude in the order of $\pm5^\circ$ or more in the order of $\pm15^\circ$?
Can this be assumed to be the same as the maximum blade flapping angle?

Assuming you mean, in-flight, it depends on the main rotor design. A fully articulated head can allow for angles great enough for an H53 to chop-off it's own refueling boom (See Here: https://www.bing.com/videos/search?q=h53+chop+fueling+boom&qpvt=h53+chop+fueling+boom&view=detail&mid=1AD065C255D15A52813F1AD065C255D15A52813F&&FORM=VRDGAR ).
Helicopter designs employing non-articulated rotor systems often have a cockpit indicator for "Mast Moment" in order to alert the pilot that excessive cyclic input is being commanded, risking Rotor/Head/Mast failure. These types require the blades themselves to absorb most or all of the rotor's dynamic articulation forces, hence much less angular deflection of the disc.
Additionally, the duration of the deflection needs to be factored-in. During maintenance test flights of the old UH-1 Huey, a "Pylon Rock Check" is performed. From an IGE hover an instantaneous full deflection and return to neutral (less than 1 sec for the full maneuver) is conducted either left or right. Somewhat violent and not something you would want to do every day for the sake of the transmission mounts. Sorry I can't provide exact disc angles though.

That's all interesting information, but it sounds like you don't know the answer to the question being asked.

I don't know the answer.This question is somewhat like asking how much control deflection does it take to snap-roll a 747.

Every rotor head has its stops, and the flapping articulation has its own stops too. The maximum deflection of the rotor disk is different according to the construction, but it's always the deflection allowed by the rotor head stops, plus 1/2 the flapping amplitude allowed by the flapping stops...

Plus the extra deflection due to the flexibility of the blades themselves, that –under extreme conditions– may be enough for the tail to be chopped off...


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Why do the inlet cones on an SR-71 retract at higher speeds while the one on a MiG-21 extends at higher speeds?
As shown by this illustration in the SR-71 flight manual, its inlet cones (spikes) retract as its speed increases:
 (source)
Meanwhile, the Wikipedia page for the MiG-21 describes the opposite behavior for its intake cone:

On early model MiG-21s, the cone has three positions. For speeds up to Mach 1.5 the cone is fully retracted to the maximum aft position. For speeds between Mach 1.5 and Mach 1.9 the cone moves to the middle position. For speeds higher than Mach 1.9 the cone moves to the maximum forward position.

No source is provided for this part of the Wikipedia article. If the info about the MiG-21 is accurate, why does its intake cone behave differently than the SR-71's? Do the two aircraft's intake cones serve different aerodynamic purposes?

You need to look at the bypass flows to understand spike scheduling. The spike moves back above Mach 2.5 to reduce throat area and limit intake mass flow. The larger inlet area doesn't match that – the excess air at the inlet flows through the bypass vents aead of the throat straight back to the afterburner. More air is sucked through the boundary layer removal vents on the cone and flows through the struts to the outside of the nacelle.

This is only a guess, so please take this with a grain of salt.
An important characteristic of supersonic engine inlets is the total pressure loss (or also sometimes pressure recovery). The total pressure loss over a normal shockwave is a lot higher than if you would slow down the air to the same speed through several oblique shock waves.
If you want to understand why the sr-71 spike retracts I recommend you watch this video. The sr-71 had a quite complex inlet geometry to ensure that the air was slowed down over many oblique shocks.
sr-71 engine inlet
On jet engines that don’t have such a complex inlet geometry as the sr-71, the best pressure recovery is mostly achieved by getting the primary oblique shockwave from the spike or inlet ramp, to hit the leading edge of the engine inlet, where it gets reflected as another shockwave.

As speed increases the primary oblique shockwave that gets formed on the spike gets shallower meaning that if you wanted to keep the shockwave of the mig-21 spike at the leading edge of the engine inlet, you need the spike to move forward as speed increases.

I suspect the difference in engine types also plays a part: the MiG-21 uses an afterburning turbojet, while at high speeds, the SR-71 uses a ramjet, with virtually all the intake air bypassing the turbojet core and going directly to the afterburner.

@Mark that shouldn't make any difference for the design of the inlet. Your two goals for both are to get the air subsonic so you can properly burn your fuel and to achieve a good pressure recovery.

Unfortunately this answer is going to be speculative as I was not part of the design team for either aircraft. However I would have chosen similar behavior.
The first thing to note, is neither of their spikes are retracting before mach 2.5 (according to the info you provided).
The mig doesn't say what it does above mach 2 but up to mach 2 they are both extending.
Extended spikes are important when traveling at super sonic speeds as your air-craft gives the air in font of it no warning that you're about to hit it.
This means most "aerodynamic" features such as round curves become almost pointless as there is no way to "slowly" push the air out of the way.
For a more detailed explanation of this you can check out this video:

Now as to why they retract at even higher mach speeds you have to remember these engines are turbo jets. Which means ideally you want combustion to happen in-between the compressor and the turbine. This happens best at sub sonic airflow, faster than that it becomes difficult for the fuel and air to mix and ignite.
That means the engines need a way to slow down and regulate the airspeed and that's exactly what the bleeder values are doing behind the spikes. The less air you let into the engine the easier it is to slow down as there is less combined momentum.
If you find more information about exactly when the spikes are retracting in the SR-71 it'd be interesting to read about. However given the diagrams above this is what I think is happening.


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Is it Legal to listen to Air Traffic Control Frequencies in the USA on a computer for entertainment?
like listening to scanners an radio frequencies for entertainment and i make it a game because i use a online flight tracker and try to find that aircraft. Is it legal?

Just to be clear, while doing this is legal in the USA, it is not necessarily legal in all countries. Having said that, whether the legality is practically enforceable or not is another question.

Yes it is legal in the US, and you can listen to audio streams on websites such as https://www.liveatc.net


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Why is there still a preference for turboprop airliners over the new regional jets?
In the 1950s and in the early 1960s, jets were mostly available as fighter/military or big, long haul aircraft. Regional aviation was based primarily on piston aircraft that could operate in smaller airports.
When planes like the Boeing 727 and 737, Douglas DC-9 and Hawker Siddeley Trident came out, they finally brought the jet era to the regional aviation. But still, they were loud and inefficient. That's where the turboprops come in: they perform better than pistons and are suitable for small airports; they don't go as fast as jets, but are usually cheaper to maintain.
Given all that, most regional airlines started to use turboprops for their city links.
In the recent years, however, jet engines are getting quieter, more efficient and powerful. Embraer's Regional Jets and E-jets are out there; more recently, we've seen Bombardier releasing their CSeries and Embraer selling the first E2 units. Even though these planes are highly efficient, fast and suitable for short runways, airlines from countries like Brazil, Indonesia and India are still using turboprops as almost the entirety of their regional fleets. The development of modern planes like the ATR72-600 and the Bombardier Q400 shows how high the interest from airlines is.
The question is, what could be the cause of such preference? Why don't airlines prefer regional jets over turboprops, given their efficiency and speed and considering the tendency people have to "fear" propeller planes as old and untrustworthy?
Speaking of numbers, what would be financially better for a company: flying an ATR/Dash 8 or a modern jet like the E2/CSeries?

The simple answer is turboprops are more profitable on short stage lengths, say up to 400 miles or so, where the limited ceiling and slower speed are less of a detriment and the turboprop's better seat mile fuel burn is a bigger factor.  This allows lower ticket prices. Once you are getting out toward 1000 miles or more, it becomes worthwhile to go higher for both efficiency and weather avoidance ability, and jets start to become more profitable.
The -8 400 extends this somewhat and can almost match jets on moderate stage lengths on total trip time, so it bridges the gap between TPs and jets to a degree.  It's noise and vibration cancellation system helps this a bit by making the trip a bit less unpleasant.
There are other complications.  Depending on the market, pax may be willing to pay more for jet service on short trips, especially in affluent markets.  There are RJ operators in the US that run CRJ200s or 900s on very short runs that would be way cheaper served by TPs because their customer base demands jets and it's worth running those routes at razor thin margins to keep the customer base.
Other factors include things like maintenance sophistication.  Older TPs require less electronics-savvy technical expertise and a small operator having great difficulty retaining mechanics trained to figure out software based control systems will be attracted to simplicity.
One of the biggest drawbacks to the -8 400, and which in truth is killing it vs its ATR counterpart in sales, is the level of electronics sophistication is high compared to other turboprops, nearly jet-like.  When Canadian regional Air Canada Jazz acquired the -8 400, their -8 technical team was warned by their RJ technical team to get ready because the airplane was closer to a CRJ900 maintenance wise (systems run by black boxes vs hydromechanical components that any wrench turner can figure out) than an older -8.  The training level of the technicians becomes critical in achieving decent dispatch reliability.
TP vs C Series (soon to be renamed the Airbus somethingorother)?  Same thing. Depends on the mission requirement.  The C Series' competition is other jets, not turboprops.

It's not to hard to demonstrate why both are needed.  Turbo props are smaller, burn fuel well under 10,000.  Speed isn't really a problem when you are only traveling a few hundred miles.  A pure Jet wouldn't even get to altitude where the air is thin representing a drop in drag. A short route would not even allow it to climb to the thin air.  In the end it's all about the benjamins.


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What are the alternatives to XFLR?
So, currently my team and I are working on a glider project for our senior design.
The design we've come up with causes XFLR to crash for some unknown reason (the issue seems to be at the root airfoil which is wedge shaped). What I would like to know is if there any viable alternatives to XFLR for airfoil and aircraft analysis that won't crash like XFLR does.
Edit: So the main goal of this project is develop an autonomous system to be used in hypersonic glide vehicles. While obviously we're not going to be making a hypersonic glider, my professor is insistent that it at least looks hypersonic. It needs to be able to glide for about 2000 feet after being dropped from a height of 300 feet.

Are you trying to characterize airfoils for use in hypersonic flight?

my professor is insistent that it at least looks hypersonic. Please tell me that this is a professor in the field of visual communication or marketing, not in the field of aerospace engineering.

What is the criteria for "looks hypersonic"? That aside, this question is about resource location, which is off-topic here.

@RalphJ basically, similar in appearance to the x-20 or x-30

@DeltaLima let me rephrase. He wants it to appear as though it has the characteristics of a hypersonic vehicle. He is an aerospace engineering professor

He's trying to use XFLR to analyze an airplane. Why is this off topic?

My initial response is that, if you have a wedge-shaped airfoil on a glider, what sort of spec are you designing to?  Most of the time, gliders are low-subsonic aircraft with airfoils similar to Epplers, Wortmann FX, or 4-series NACAs, not wedges.  I associate wedges with massive flow separation or supersonic flows.  Can you provide a picture of what you're talking about, as well as a discussion of what you're designing to?  Also, are you attempting to produce a 3D analysis of a wing that, at its wing box, has this wedge-shaped airfoil?
That said, my short answer is this: XFLR is a panel-method solver, which means that it cannot deal with flow separation save by approximation.  As a result, shapes such as a wedge, which are quite prone to flow separation by virtue of their shape, are very hard, if not impossible, to analyze with these solvers.  Unfortunately, lower-order solvers (i.e., something short of finite-element or finite-volume methods, such as Fluent, CFX, etc.) bank on using these panel methods or other lower-order fluid flow equations that cannot handle flow separation.
If you want to try to accurately model flow separation, you'll need a higher order CFD solver, which brings with it additional complexity, cost, and headache--and getting accurate drag information is a tricky prospect.  There are some freeware solvers out there, but, by far, the most popular (and widely validated) is OpenFOAM.  This is a command-line program at heart, but there are some GUIs developed for it that you may be able to obtain for free or at a reduced cost, such as SimScale.  Some other packages are discussed here as well, if you'd like to have a look.
That said, what data do you need?  If all you need is a sectional drag coefficient, I would highly recommend a look through Hoerner's Fluid Dynamic Drag (not to be confused with his other text, Fluid Dynamic Lift), a vital reference that contains data from a plethora of experiments done on a multitude of shapes.  In it, he describes the drag force on objects from ellipsoids to cones to cylinders and more, all based on size, aspect, roughness, and Reynold's number.  There is an updated copy available, but I haven't found it online as of yet.

So the main goal of this project is develop an autonomous system to be used in hypersonic glide vehicles. While obviously we're not going to be making a hypersonic glider, my professor is insistent that it at least looks hypersonic. It needs to be able to glide for about 2000 feet after being dropped from a height of 300 feet.

@ironduke97 That's a pretty important bit you should probably mention in your OP...what flow regime are you running in the solver?

@Jimmy so far, we've run analysis for Reynolds numbers from 100k to 900k in increments of 10k. The glider is only travelling at around 22 mph or so.

This just got more interesting.  Yes, XFLR is definitely not the tool you want.  There is some evidence for hypersonics done well in OpenFOAM and variants (https://bit.ly/2maCEwl), but it is going to be much more difficult.  Possible answers of the non-CFD variety: https://aviation.stackexchange.com/questions/31607/what-kind-of-airfoil-configuration-is-preferred-for-an-hyper-sonic-plane, https://thesis.library.caltech.edu/5025/1/Wunch_ws_1948.pdf.

From what I've gathered, you are designing a small, low speed, unpowered aircraft with a glide ratio of 6.7 that must look like a hypersonic vehicle. In order to meet the latter requirement, you chose a wedge airfoil. I'm also going to assume that your aircraft has very low aspect ratio, highly swept, thin body that generates most of the lift.
I wouldn't worry too much about the airfoil. Just assume it's a flat plate. Try it out with XFLR's panel method. If it still fails for some reason, then just go with a VLM method like AVL. It should get you into the ball park for lift. Maybe assume a $C_{L_{max}}$ of 1 for conservatism.
For drag, just use the Blasius solution for skin coefficient; or you can use Hoerner's as suggested by the previous answer. Make sure you have the right Reynold's number. 900,000 as you mentioned in the comments seems way too high for what you have described.

You pretty much got it right. The only issue is a lot of the points are unconverged and we get a lot of "outside flight envelope" errors

@ironduke97 Have you tried AVL?

I've only heard of it


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Sensitivity of aircraft stability with respect to the CG location
I am working on the design of an aircraft (2 m long, wingspan = 3 m). The platform is longitudinally stable but it is very sensitive to the variation in CG location. For example,
1) CG_location = 100 mm from wing L.E, the static margin is about 15% of MAC.
2) CG_location = 105 mm from wing L.E, the static margin is about 8% of MAC.
3) CG_location = 110 mm from wing L.E, the static margin is about -1% of MAC. (Unstable: Cm vs Alpha curve slope becomes positive)
So, effectively while manufacturing the aircraft or while dropping the payload the CG of my aircraft should not vary more than 1 cm, otherwise it would become completely unstable. Taking into any manufacturing errors, the platform CG might not be exactly as expected and hence having such a small margin creates issues.
How should I deal with this issue? Can I have any suggestions on how do decrease the sensitivity of the longitudinal stability to the CG location? Thanks in advance.

It sounds like it has a horizontal tail, so you basically need to increase the tail volume which will allow you to move forward CG limit forward, allowing you to move the median CG forward and expanding the overall range. So just make the horizontal tail a bit bigger.

Or increase the tail arm.

Tail volume is area times arm.  You can increase one, both or all three. Making the surface bigger is easiest.

Using your numbers, it looks like your MAC is about 71.4mm and NP is about 110.7mm aft of LE. That means a 1cm shift in CG is equivalent to 14% MAC, which sounds about right. That is, longitudinal stability is not sensitive to CG location.
For a small drone like this, keep in mind that CG isn't as sensitive to manufacturing as you may believe, unless you have a superb payload to weight ratio. If that's the case, you can increase the tail volume as others suggested to shift the NP aft (thereby increasing the static margin).

10 mm CG shift is 16% MAC in the OP's numbers. That makes MAC = 62.5 mm. This is very narrow for a 3 m wingspan (aspect ratio = 48). But if this is true, then indeed, there is nothing particularly wrong. With the NP already almost a chord aft of the wing, the tail volume must already be quite big. With the arm about 1.5 m (of the 2 m length), damping will be enormous (and downwash small), so I bet the model will be flyable even with very low static stability.

Thanks for catching this. Yes, this is indeed an enormous aspect ratio...(larger than the Rutan Voyager). Is this intended by the OP?

That's the question... I don't know. It's not impossible, but it's significantly more than the champion-level gliders of comparable size.

A picture of the design would be helpful, and also a caveat on the legality of "dropping something from a plane".
Now, in addition to what John K suggested, a real quick fix would be to lengthen the fuselage
a full meter aft and fore as well.  This increases the torque of your tail (I would make it bigger too).  This not only gives you a wider CG range, it also slows down the pitching movement, making the plane more controllable.  Particularly if extreme range is not absolutely necessary, go ahead.  Lowering the aspect ratio of the wings (making them wider)
also will help.
As a point of interest, many people are flying quad rotors around these days with possible
plans for package delivery.  Like a miniature Chinook helicopter,  they can tolerate a very wide range of CG fore and aft simply by adjusting power to the rotors.
But for an aircraft, look to the wing, the fuselage, the tail, or all.  Any payload, just like real cargo carriers, needs to be centered at CG as much as possible.  A moving or sliding adjustable cargo floor may help your design.

Really, he has the choice of making the tail boom longer, the tail larger, or both. Any one will increase tail volume (area x arm).


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Contribution of wing vertical location to roll stability
Recently, I was reading about the factors that contribute to roll stability of an aircraft and came across a point stating that the wing's vertical location (Distance of wing from CG) contributes to the roll stability of the aircraft. I cannot comprehend this statement completely and even if it did, How much of an effect does it truly cause when compared to having features like a wing dihedral? Any information and explanations would really help me understand this concept.

Possible duplicate of What are the dynamic stability properties of an aircraft with wings exactly at the vertical CoG?

The Center of Gravity of aircraft is not a vertical measurement, it is a horizontal measurement from the datum point of the aircraft, or expressed as a percentage of the mean aerodynamic chord. High wing or low wing is not a contributing factor to roll stability, but the wing's dihedral angle is.

@JuanJimenez  There is a vertical CG, although until you have the height of a 747 it's typically ignored during load planning. But it's there, and designers are aware of it, even if crews don't have to be (given how little they can affect it).

Voting to leave open, since nothing in the other answer really addresses the OP's question other than to say "aerodynamics matter more". Okay, that's great, but these effects seem worth discussing, even if they're less influential than the other.

@JuanJimenez As far as physics and everyday usage are concerned, an object's centre of gravity is a point.

Related: Why do low wing aircraft have higher dihedral than high wing aircraft?

@JuanJimenez The c.g. is defined for all 3 axes, although the longitudinal c.g. is the only one that is generally checked for all aircraft. However, wide-bodies flying cargo typically check the lateral c.g. by means of making sure they don't exceed a lateral moment limitation, and 767s have vertical c.g. limitations. At jwab.website you can play with 747 and 767 cargo c.g.s if you want.

An aeroplane in a pure co-ordinated turn has a deflection of both the ailerons and the rudder. This is cumbersome however, much easier to only deflect the wheel, tilt the lift vector, and let the aircraft sway into the turn. The vertical tail will then experience a slip angle, which will point the nose into the direction of the turn at the expense of a small sway velocity: sideways.
So now the question is: how does the aeroplane react to a sideways incoming airstream. What we want is that if the pilot deflected the wheel to the right, he has to keep the wheel to the right to maintain the right hand turn. So upon an un-coordinated turn, without touching the pedals, we want the aircraft to try to level itself again upon an incoming sideways velocity.

If we look at what happens to a low wing aircraft first, we see that the incoming sideways stream sees little resistance when streaming over the fuselage top, but gets caught in the wing/fuselage intersection and builds up pressure there. So the sideways velocity introduces a rolling moment in the direction of the turn, which the pilot has to compensate for by turning the wheel the other way. This is not the behaviour that we want.

In order to correct for that, we can introduce wing dihedral: now the incoming sideways air stream wants to level the aeroplane again, and the pilot needs to keep the wheel deflected into the turn, which is what we want.
And of course, with a high-wing aeroplane the above is valid vice-versa: the sideways stream wants to level the aircraft, so the roll behaviour is naturally how we want it to be already. Depending on the wing/fusleage shape and many other factors, the aeroplane can be too stable and we'll need to implement anhedral, wings pointing down.

Draw a vertical line through the center of lift in the undihedralled low wing at a bank and you will see the displacement of CG from Clift.  In a large aircraft this is what causes the roll to continue before side slip even starts.  -20, I Don't care.  This is why top heavy is unstable.

Lift goes through the CoG though. At a bank angle, lift is not vertical. And dihedral fixes roll instability...

With respect, vertical lift vectors hold the plane up. Yes, total lift and CG are still "centered" WRT "top" and "bottom" of aircraft, but they are displaced WRT gravitational field in a roll. At 500 mph this matters little, but should not be forgotten.  Displacement of lift (dihedral) or weight (hang glider) becomes important as one slows down, and as Juan pointed out, is critical in spins.  And please don't forget what slip does to a tall upright vertical tail.  Regards, Bob

High-wing airplanes are usually designed with less dihedral than low-wing airplanes because of the various different ways that a high wing placement enhances roll stability.
Several of these are addressed in this answer-- Why do low wing aircraft have higher dihedral than high wing aircraft?
For any of this to make any sense, it is imperative that we understand that when an aircraft is banked, it tends to turn (i.e. the flight path tends to curve), and in a turn, there tends to be sideslip, unless the pilot is using the rudder to eliminate it.  Sideslip is the fundamental driver behind nearly all the dynamics at play with regard to roll stability.  The actual reasons for the sideslip are not obvious and are related to the fact that in a steady turn, the wingtips are moving at different airspeeds, which tends to create a yaw torque.  Also, if the vertical fin is tangent to the curving flight path through the airmass, then the fuselage centerline at a more forward location on the aircraft cannot be tangent to the curving flight path through the airmass, so there is no way for the fin to completely eliminate sideslip in the absence of a rudder input.  In fact, in some cases the vertical fin can be said to actually be causing sideslip.  (Imagine a very long fuselage and a small radius of curvature of the flight path.)
Therefore a non-level bank attitude causes a turn which causes sideslip which can create a roll torque towards wings-level via dihedral, "pendulum effect", and other aerodynamic effects.
Note that in a high-wing aircraft, the wing's drag vector, acting above the CG of the aircraft, will tend to contribute a stabilizing roll torque (tending to roll the aircraft toward wings-level) in a sideslip even in the absence of any significant aerodynamic sideforce.  This is an example of what is often called a "pendulum effect".
The "pendulum effect" is tremendously strong in paragliders, and plays a significant role in hang gliders as well.  In hang gliders, when the pilot exerts no muscle force, the pilot's weight acts as if it is located where the flexible hang strap connects to the aircraft structure.  Raising or lowering this point of attachment affects the aircraft's pitch and roll stability dynamics.  Some older hang gliders had the "hang point" located several feet below the wing, creating a strong stabilizing "pendulum effect".
Consider the difference in flight characteristics of a model airplane with a heavy weight attached to the end of a thin strut projecting high above the aircraft, versus a model airplane with a heavy weight attached to the end of a thin strut projecting far below the aircraft.  It's clear the flight characteristics would not be identical in each case.  It's also clear that if both aircraft were trimmed for a steady glide, the aircraft would be much more stable in both pitch and roll in the latter case than in the former.
See this answer for a more extended discussion of hang glider dynamics and other issues related to the "pendulum effect" and roll stability: Does "pendulum effect" apply to hang gliders or any aircraft?

"the wing's drag vector, acting above the CG of the aircraft, will tend to contribute a stabilizing roll torque" – that you need to explain. Drag points backwards (per definition). How can it produce a roll? Due to its backward orientation, it should only be able to cause yawing and pitching moments. What would be the effect for causing a rolling moment? That would require either sideways or vertical drag. Those forces are, however, called side force and lift, respectively.

@PeterKämpf if we define roll as motion around the aircraft's longitudinal axis, then when the aircraft is yawed to point in a different direction than it is moving through the air, a drag force acting above or below the CG will create a roll torque.

So this is a sideslip-induced rolling moment. It has several contributors, and drag is only a small part of it. It is dominated by dihedral. And a pendulum is really not involved.

Lovin' the dv's as always.  This was actually one of my better answers ever.  Short, to the point, and only lightly edited!


text:
Propeller effect on the L/D of an aircraft (Pusher Propeller configuration)
I recently performed an aerodynamic analysis of a fixed-wing UAV with a pusher propeller configuration. I found out that the L/D reduced by about 15% when compared to the body alone without propeller. I am investigating methods to mitigate this problem.
The following are some details about the aircraft,
Wing Span = 2.5 m
Fuselage length = 2 m
MTOW = 25 kgs
Tandem wing configuration. The following image is a representative image.
Courtesy: Raytheon Coyote

Any suggestions will really be helpful.

With a spinning propeller, or propeller at standstill?

Please explain how you measured that L/D reduction. Was the propeller spinning or fixed? How did the end of the aircraft look like initially, without the prop?

You need to fair in the tail with a carbon fibre fairing/spinner such that the prop folds back and nests into it.  Like below.

Thanks for the reply @John K. I have incorporating such a fairing mechanism at the rear portion and the above problem I mentioned is still persistent. Are there any other methods to decrease the negative impact of the propeller?

The only other thing that might help is the extend the spinner in a way that lets the blades lie completely flush in the spinner profile when folded.

Noted. Will try to push that to the greatest extent possible.


text:
Aerodynamic Pressure Loads in FEA of Lifting Surfaces - Query
Let me give a brief summary of the process involved in the FEA of lifting surfaces with the wing as an example:
1. The wing is modelled as a hollow surface body in Ansys, meaning only the exterior aerodynamic surface is specified.
2. Internal structure made of composites is modelled next.
3. Aerodynamics loads and Boundary conditions (fixed edge at the root of the wing) are specified.
4. The solution is obtained in terms of displacements, stresses and failure criteria etc.
As part of Step-3, aerodynamic loads are generated in the form of Cp values (Coefficient of Pressure) at discrete points on the exterior surfaces of the wing. This data is generated from XFLR5.
Cp values are converted into pressure values and applied on the exterior surfaces of the wing. To do the same, the following formula is used:

Pressure values obtained from the LHS of the above expression are applied over the exterior surfaces of the wing. FEA process is completed and the solution is obtained.
FEA predicts that the wing cross-section is unable to maintain its airfoil shape and the wing cross-section essentially gets "crushed"'.

Observe the crushing of the wing cross section in this figure where the top surface and bottom surface come close to each other and airfoil geometry is completely lost
However, it finally struck me that there has to be an internal air pressure that will support the cross section from the inside as what happens with hollow structures on earth, eg. a hollow can be placed on the ground doesn't get crushed into itself since the external and internal pressures cancel each other out.

Hence, we should not be using absolute pressure but rather gauge pressure (difference in external and internal pressure) should be used. Otherwise, it essentially means that there is a vacuum on the inside and air pressure on the outside.

Once this change was made, we no longer observe the "crushing" effect.
Finally, we still want to be sure if the internal air pressure would be equal to the static pressure at that altitude as has been assumed in the usage of gauge pressure.
Our assumption is that the pressure inside the wing body would be equal to the static air pressure at that altitude from thermodynamic relations even if the wing is completely sealed. IS THIS ASSUMPTION CORRECT? ANY SUGGESTIONS WOULD BE HELPFUL?

If your question is "is pressure inside the airfoil equal to outside static pressure?" you should edit the title to reflect it.

There is fuel inside the wing, and there is definitely a connection with the outside atmosphere for venting while fuel is flowing into the engines. The fuel tanks are sealed from the outside airflow. In a document about the B737NG from Smartcockpit.com (now via Scribd) it says:

The purpose of the fuel vent system is to prevent damage to the wings due to excessive buildup of positive or negative pressure inside the fuel tanks and to provide ram air pressure within the tanks. The tanks are vented into surge tanks which vent through a single opening at each wing tip.

A recent development, long recommended by the NTSB but deemed too expensive to mandate by the FAA, from this site:

Boeing has developed a Nitrogen Generating System (NGS) which decreases the flammability exposure of the center wing tank to a level equivalent to or less than the main wing tanks. The NGS is an onboard inert gas system that uses an air separation module (ASM) to separate oxygen and nitrogen from the air. After the two components of the air are separated, the nitrogen enriched air (NEA) is supplied to the center wing tank and the oxygen enriched air (OEA) is vented overboard. NEA is produced in sufficient quantities, during most conditions, to decrease the oxygen content to a level where the air volume ... will not support combustion. The FAA Technical Center has determined that an oxygen level of 12% is sufficient to prevent ignition,

I wouldn’t work out the static pressure at the inside as equal to the static pressure of the undisturbed air away from the plane because the low pressure is forced at the inside of wing due to outside airflow vacuuming the inside air. What you need is to build an internal structure, add resistance to the skin and redo the calculations. Let the inside static pressure be as low as it should be as a result of airflow, as if the inside communicated directly with the outside at trailing edge and a differential vacuum ( pressure lower than the outside static air pressure) was created internally. I would do that based on the fact that control surfaces ( flaps, ailerons) allow free air flow between inside and outside.


text:
Do floatplanes land on rivers?
I'm sitting here in a coffee shop on the local river and I noticed it was just wide enough to facilitate a landing. There is no chop on the river, but it does move pretty switfly, and there are bridges and whatnot that would quickly come into play. However, there are some pretty slow-moving, unobstructed rivers out there.
Does anyone land floatplanes on rivers for reasons other than emergencies?

Welcome to aviation.SE! You might find this answer interesting

Next question, do they land on canals?

Philosophical question...is putting down on water actually "landing"? I know it's not watering!

@JohnK please post answers in the answer field.

@Frederico the question was already answered several times I was just adding additional info.  Which is what comments are for no?

Float planes are used world wide on rivers. In Alaska and Canada the primary use is fishing, hunting, sightseeing, lodging, supplies...

Cessna 206 delivering supplies to a fishing camp along the Yukon River (Alaska).

Sightseeing along the Amazon River.

Wind River (Canada) tour map

Yes, here is a river landing.

and another, even tighter. This might technically be a beaver dam pond in a river.

and another, prettier. I'd back this one up and watch the approach.

They absolutely can and do on a regular basis.  River currents will play a significant role in takeoff and landing considerations as well as during displacement taxiing, casting off, docking and mooring of the airplane.  But it does happen on a regular basis.


text:
How much does the horizontal stabilizer on a 737 move for one unit of trim?
Inspired by this question, what exactly is the meaning of a stab trim unit? Is it exactly one degree of stabilizer movement?
The 737 NG FCOMv2 (9.20.9 Flight Controls - System Description) only lists the ranges of the units:

Stabilizer Position Indication and Green Band
Stabilizer position is displayed in units on two STAB TRIM indicators located
  inboard of each stabilizer trim wheel. The STAB TRIM indicators also display the
  TAKEOFF green band indication.
The trim authority for each mode of trim is limited to:

Main Electric Trim

flaps extended 0.05 to 14.5 units
[737-600] flaps retracted 4.10 to 14.5 units
[737-700] flaps retracted 4.30 to 14.5 units
[737-800] flaps retracted 3.95 to 14.5 units
[737-900] flaps retracted 3.90 to 14.5 units

Autopilot Trim 0.05 to 14.5 units
Manual Trim -0.20 to 16.9 units.

The green band range of the STAB TRIM indicator shows the takeoff trim range.
  An intermittent horn sounds if takeoff is attempted with the stabilizer trim outside
  the takeoff trim range.

The full range of travel can be seen as a black arc in the following image (taken from this answer), which could be 17°, but it is hard to say from the image:

The stab trim unit is an offset to the actual horizontal tailplane angle. It is added by Boeing to make the trim value positive.
1STU = 1 degree

Do you have a source for that?

This post on pprune.org confirms your statement. It also does not list any sources, but I'm going to accept for now...


text:
Why is the autopilot of the DHC-6 Twin Otter powered by bleed air?
I was just reading through the POH of a DHC-6-100 Twin Otter and was surprised to read the following:

3.8.2 LOW PRESSURE - OTHER SYSTEMS. The airframe deicing system and
  the autopilot are operated by bleed pressure. In the event of a low pressure
  condition, a caution light on the caution lights panel will illuminate. Under these
  conditions, the pressure may be restored as outlined in paragraph 3.8.1. However,
  if increasing engine speed does not rectify the condition, the use of the
  autopilot and de-icer system should be discontinued.

(emphasis mine)
I understand why the deicing system is using bleed air, but why is the autopilot powered by bleed air? How does this autopilot work?
The following question is somewhat related:

How did old autopilots work?

but it never mentions anything powered by bleed air.


text:
Can a Bombardier CRJ derate the takeoff thrust?
I have recently been flying a Bombardier CRJ 700 in a simulator and wondered how to derate the engines for takeoff.
I know how the takeoff thrust derating is selected in Airbus and modern Boeing aircraft via the (M)CDUs (see also How do pilots set the exact amount of thrust needed for a reduced power takeoff?) and I have also seen panels for setting engine modes including a flex temperature in aircraft like the MD-80 or the Avro RJ. However, I could not find anything like this on the CRJ (at least not in the simulator).
Can the CRJ even derate the thrust for takeoff? And if, how does it work?
I saw the thrust levers have detents labelled SHUT OFF, IDLE, CLIMB, TOGA, MAX POWER:

(image source)
I assume TOGA would result in less thrust than MAX, but how is this selected?


text:
What aircraft is shown in this cockpit picture?
While watching this YouTube video, I saw a stock image of a cockpit (at 28:55; unrelated to the rest of the video), which I didn't recognise:

The general design and some panels look very Airbus like, but it doesn't agree with any Airbus model I'm familiar with. In particular, the screen layout is quite different.
A Google reverse image search brought me to this Shutterstock image, which shows a wider angle (and also the typical Airbus fire handles), but the description doesn't help:

Flight Deck of modern aircraft. Pilots at work. Clouds view from the plane cockpit.

Which aircraft is shown in the image?

Yes, youtube creators make extensive use of stock images.  You found the exact image.

I am fairly certain what you're looking at is the cockpit of an SSJ100, it has a similar sidestick layout to Airbus, but it was the window opening mechanism that is quite distinctive.

The layout of controls and screens also very closely matches your original picture, as well as the blank spaces where a clipboard exists in the wikimedia picture, and what looks like a tablet computer exists in yours.

https://www.airliners.net/photo/Interjet/Sukhoi-SSJ-100-95B-Superjet/2416045/L

Yep, that's definitely it. I was wondering why it looks so Airbus like. The Wikipedia page says "Key suppliers were selected in October 2003, including Thales for avionics [...]" and Thales is also providing flight deck avionics equipment for Airbus.


text:
What RNP level is required for an RNP approach if the chart does not specify a value?
RNP (Required Navigational Performance) approaches require the aircraft to monitor the navigation accuracy
within a specified limit, called RNP level:

There are several different levels of RNP. Examples of RNP levels used for approach include RNP 0.1, RNP 0.3, and RNP 1.0 (There are also RNP 4.0 and RNP 10.0 levels that apply in the en route environment). "A performance value of RNP 0.3, for example, assures that the aircraft has the capability of remaining within 0.3 of a nautical mile to the right or left side of the centerline 95 percent of the time."

[...]
The minima line will include a performance value, RNP 0.30 for example. GPS with WAAS can support RNP 0.3 and above.

(faa.gov on Satellite Navigation)
Usually the RNP level is given in the minima section of the approach chart (and according to the above FAA quote, this is always the case in the US).
There are however approach charts that do not include this. Consider for example the following 3 approaches into Innsbruck (LOWI), Austria. While the first example lists the RNP levels, the other two do not:

RNP Z RWY 26 (AR)

There are two different minima for the different specified RNP levels.

RNP E RWY 26 (LPV ONLY)

There is no RNP level specified, also not anywhere else on the chart (as far as I could find). Maybe this is because this approach is LPV only (requires EGNOS1)?

RNP Y RWY 08

Again, there is no RNP level specified and this time there are only LNAV minima. So which RNP level would apply here?

1 EGNOS is the European SBAS (Satellite Based Augmentation System), equivalent to the North American WAAS.

RNP approaches require a minimum value of 0.30, be it LNAV only, LNAV/ VNAV or RNP AR. Most of the approaches around the globe are RNP approaches. So, they do not put in the value because it remains the same at 0.30.
In RNP AR approaches they put in the value because these are approaches that require specially trained pilots and specially equipped aircraft. That is where the AR comes from. It stands for Authorization required. The RNP AR approaches have RNP values starting from 0.30 down to 0.10. The biggest difference between a typical RNP approach and an RNP AR approach is the lateral obstacle clearance. In an RNP approach, you have a clearance of two times the RNP value (0.30 x 2 = 0.6 miles) and also a buffer of 0.2 miles either side. In an RNP AR approach, the buffer is taken out of the picture and it only offers two times the RNP number left or right. If the RNP value is 0.10, then the lateral obstacle clearance will be 0.20 miles either side. This allows tighter approaches in areas with high obstacles such as mountains (such as Innsbruck). But because such approaches are a bit demanding you need better equipment. So, airplanes that are not designed for it need extra modifications. These modifications will depend on the manufacturer and it will also depend on the actual procedure. The pilots who fly these approaches need extra training as well, so that they are more accustomed to flying it and to teach them the required contingency procedures. Depending on the operator and the aircraft configuration, an RNP AR approach maybe approved with an RNP value between 0.30 and down to 0.10. For example, Airbus offers A320 and A330 with modifications and approvals for either RNP AR approaches with 0.30 or 0.10. In case of the A350, the aircraft comes with RNP AR 0.10 approval as a basic setting.

For LPV approaches, there is not an RNP value as it is basically an ILS type approach flown to category I minima using the GPS and a ground based GPS signal augmentation system such as WAAS (Wide area augmentation system).

@assylias, thanks. It is corrected now.

All those approaches require RNP APCH capability, i.e. a precision of 0.3 on the final segment (see table 1 under GM1 SPA.PBN.100 of Annex V
Specific approvals via easa.europa.eu):

They also have a RNP AR APCH option (which requires additional equipment and a specific authorisation), with a 0.15 precision requirement. I suspect that indicating 0.30 on the "standard" Z approach is to avoid any confusion and clarify that usual RNP APCH equipment is appropriate.


text:
Why are there large gaps in the spoilers of the Airbus A300?
The spoilers on the Airbus A300 have large gaps at the bottom:

(A300-600R, modified from Wikimedia Commons, original picture by Tennen-Gas, CC BY-SA 3.0)
I have never seen any other aircraft with such gaps in the spoilers. Airbus also got rid of them on the A310:

(A310-322, ETH-Bibliothek Zürich, picture by Swissair, CC BY-SA 4.0)
Since the point of spoilers is to disrupt airflow on the top of the wing to reduce lift, it seems counter-productive to include such gaps. So what is their purpose?

I cannot answer with scientific accuracy, but the gaps do not diminish the efficiency of the spoilers in a significant manner, but relative to their size the pitching moment and actuation force required are smaller compared to a "solid" version.

What do dive brakes have to do with it????


text:
Why is the LOC RWY 26 approach at Innsbruck classified as LOC instead of ILS or IGS?
After reading about Localiser Type Directional Aid (LDA) and Instrument Guidance System (IGS) on SKYbrary, my understanding of classifying these approaches is this:

Course Offset
With GP
Without GP

up to 5°:
ILS approach
LOC approach

more than 5°:
IGS approach
LDA approach

(In the US (TERPS), the limit is only 3° course offset, but ICAO Doc 8168 allows up to 5°)
However, the LOC R Rwy 26 approach at Innsbruck (LOWI) is classified as LOC, even though it has a GP (and a course offset of 3°). AFAIK, the approach chart does not indicate at all that a GP is available (correct me, if I'm wrong):

But when digging into the AIP, one finds the GP information:

(Austrian eAIP: LOWI)
The GP frequency of 331.7 MHz is paired to the LOC frequency of 111.1 MHz (example source). So this LOC approach has a paired GP. Then, why is it not classified as ILS?
I suspect that it might have something to do with the LOC antenna location, which is not on the runway centerline:

(Austrian eAIP: LOWI ground chart)
However, if this offset no longer satisfies the conditions for an ILS, then shouldn't the approach be classified as IGS instead (because it has a GP)?

Having transmitters for lateral and vertical guidance doesn't mean they satisfy the required tolerances for an ILS/IGS, as defined in ICAO Annex 10, volume 1 (3.1.3 and 3.1.5). An example can be found in: What are the reasons for autopilot restrictions on instrument approaches?
In the case of LOWI, the likely reason for the glidepath (GP) navaid not meeting the standard is the approach path goes down in the steep-sided Inn valley:

LOWI airport, seen from West, source Shutterstock
This might be due to reflection on the mountains around (there are also other suggestions like temperature and snow affecting the radio signal).
The GP guidance is thus provided only for the purpose of monitoring, but it can't be used to follow the glidepath with the autopilot. Pilots must use the DME and the altimeter. Extract from AIP:

A localizer (OEV 111,10 MHZ, LOC course 255° MAG) on the aerodrome but 3,5° offset from the runway centre line and a collocated DME are providing course guidance and distance information during the instrument phase of the descent procedure and in case of a missed approach. A glidepath which is frequency-paired with LOC OEV is available coinciding with the DME descent gradient of 3,77° between D-19 OEV and the threshold.

@Bianfable. Yes, I misread your comment. I'm not knowledgeable in SOPs, but I guess as long as they double check with DME/Altitude it's ok.

Can someone expand what GP stands for?

@TomMcW GP is GlidePath, also called glide slope (G/S).

GlideSlope (G/S) refers specifically to ILS style vertical guidance from an ILS GlideSlope transmitter. It will display the GS diamond on the vertical G/S scale and the Autopilot lateral mode can be coupled using  LOC or ILS mode. GlidePath refers to a Final Approach path constructed by the FMC and flown using VNAV or equivalent FMC  mode. An ILS can have a GlidePath which the Autopilot can couple using VNAV, but a GP can't always have a G/S. A point to notice is that the grey + white LOC symbol is not depicted on the 'plan' mode chart, and neither is the GS symbol used in the profile diagram.

@skipper44: I used GP because this is the wording of ICAO related to ILS in annex 10. GP and GS are used interchangeably to refer to the transmitters and beams used to create the guidance signals. Strictly, GS is the slope of the glidepath, when the GP is a straight segment.

Though you're right as regds the interchangeable use of the terms G/S and G/P, sticking to in-cockpit nomenclature - Boeing FMA is [SPD | LOC | G/S] only if the actual radio based NAV AID transmission is present, otherwise [SPD |FAC | G/P] where FAC and G/P are calculated by the FMC. Mind you, one could also fly [SPD | LOC | G/P] or[ SPD | LOC | VNAV] but at certain airfields, Paris, CDG comes to mind,  the lowest/best DH etc. are based on an autopilot coupled approach requiring an actual ILS G/S.  I request a current Airbus pilot for their modes and FMAs.

@skipper44 The respective Airbus modes are called "F-G/S" (short for FMS glideslope) instead of "G/P" and "FINAL" instead of "VNAV" (when used during final approach). See this answer and this answer.

There could be a few reasons for no ILS Approach being published despite  the existence of GP:

steep GP of 3.77deg
GP not calibrated/'calibratable'.
NO operational benefit (eg ILS minimums worse than LOC/VOR/RNAV minimums due terrain for different obstacle surface). VOCL in India had an ILS 28 DA of around 900ft compared to VOR MDA of 700 odd ft. This is no longer true, I'm not sure of the technicalities that allowed the change, but for many years it was.
An ILS can result in an autoland and this is taboo for an offset LOC. An approach flown in IAN/VNAV (RNAV/VOR/LOC) has no autoland capability and would alert/warn the pilot very clearly if the Autopilot was still engaged as the airplane approached the RW threshold.

(Note that there do exist ILS approaches with offset-LOC, KJFK ILS 22R comes to mind.)

Thanks for the additional info! I don't think the steepness is the reason though (London City has 5.5° and is classified as ILS). Autoland only applies to Cat II or III, so that should not restrict Cat I ILS either. About ILS 22R at JFK: is that only an offset course ILS (of which there are many in the world), or actually an antenna offset from the runway centerline (I cannot tell from the chart).

I've listed factors that might prevent the approach being classified as an ILS, one, or more, or a combination could be limiting. Besides,  there's no restriction preventing an autoland during CAT1 or even CAVOK conditions - not having an ILS approach removes an opportunity for something to go wrong. The offset course is achieved by having the transmitter antenna off centre line/off RWY.


text:
Are there any autoland capable aircraft without autothrottles?
While reading the question Why couldn't the CRJ100/200 be equipped with triple-string autoland? I thought the CRJ cannot have autoland because it is not equipped with autothrottles. Somehow I always thought this would be a requirement. However, after doing some research I found that this is not the case:

(a) An automatic landing system must include automatic control of throttles to touchdown unless
it can be shown that:

Aeroplane speed can be controlled manually without an excessive workload in conditions
for which the system is to be certificated;
With manual control of throttles the touchdown performance limits of CS-AWO 131(c)
are achieved; and
The touchdown performance is not critically affected by reasonable errors in speed
control.

(EASA Easy Access Rules for All Weather Operations (CS-AWO) 123 Automatic throttle control)
This made me wonder: Is there any aircraft that does not have autothrottles or autothrust, which is certified for autoland?
Note that by autoland, I mean a proper Cat II/III ILS autoland, not something like the Garmin Autoland for GA aircraft.

You mean a system where the crew has to work the power while the airplane lands itself?

@JohnK Yes. According to the above quote from EASA, such a system should be legal. I'm just not sure if it has ever been done. All aircraft with autoland I know of, also have an autothrottle (or autothrust for Airbus).

Never heard of that kind of system, but you'd think somebody would have done it.

Isn't the general principle of landing to control height with power and speed with attitude? So adjusting the throttles while an autolander did the landing would actually be a lot of work.

@DJClayworth In general, yes, because it is safer. But not during an autoland: the A/P will use pitch to maintain the glideslope and the A/T will target the approach speed (usually Vref + 5kt).

That still sounds like a lot of throttle work. Also see this question: https://aviation.stackexchange.com/questions/81272/why-so-much-throttle-work-on-navy-fighter-during-final-approach

@DJClayworth On an ILS you don't fly the "normal" way. Because the priority is to maintain a precise slope, you "pitch to path", following up with trim and thrust changes to maintain speed.If the autopilot is doing it, the autopilot is pitching to path and trimming. On an ILS with AP on down to minimums, you just sit there making small thrust tweaks to maintain reference speed with the control column dancing in front of you.If you had an autoland system with manual thrust control, you would need a thrust reduction cue of some kind to tell you to ease off the power as autoland starts to flare.

I'll try to answer, even if it may not be the answer you expect.
To directly answer your question, yes, there are (were) aircraft with autoland system and no autothrottle. Hawker Siddeley Trident 1 is one example: it was the first aircraft to perform autoland (and subsequent versions reached also CAT II) but with no autothrottle.
However, I do not think that the existence of these aircarft is the reason for the rule. Instead, this is probably to allow the use of autoland even if the auto throttle is not working. This is, for instance, the procedures when auto throttle is not functioning in the Minimum Equipment List (MEL) of an A350:

As you can see, autoland is still available (LAND2) with limited capabilities (cannot land in CAT III). Thus, the regulation may be there to provide the conditions for the definition of MELs in case of autothrottle inoperative.
Side note: the same is impliticly stated in FAA AC 120-118, where auto throttle is required only for CAT III autoland.


text:
Why should autofeather only be used for take off on the Dash-8 Q400?
The Bombardier Dash-8 Q400 FCOM says (emphasis mine):

The autofeather system provides automatically initiated propeller feathering following an engine failure during take-off, and power uptrim of the operating engine (Figure 12.22-5). Autofeather is selected on for take off only, using the AUTOFEATHER switchlight on the engine instrument panel.

(source: smartcockpit.com)
It only says autofeather should be used for take off only, but fails to explain why. I would assume autofeather would also be useful when landing. The uptrim (automatic power increase for remaining engine) might not, but it is independent anyway:

Uptrim is triggered (regardless of autofeather selection) [...]

Why should autofeather not be selected during landing?

It's because the critical case that autofeather caters to is engine failure right after V1, on takeoff, where the normal identify/feather drill uses up precious time.  An engine failure on a missed approach or go-around starts well above the low energy state of an engine failure after V1 and there is much more time and energy margin to do the normal feathering procedure.
The system is quite sensitive and there are risks associated with having the autofeather system armed and "ready to fire" you might say.  Mainly, having it put a prop in feather when you don't want it to.  Or even BOTH props in feather.  That risk outweighs the benefit of having it armed outside of the take-off case.
It's a bit like having a policy of taking a firearm off SAFE only when there is a critical need to able to fire instantly when clearing a house.

Great answer for turboprops in general, but is there any reason why some aircraft (e.g. King Air) recommend arming autofeather for landing, and some don't?

Afraid not.  It could be that the go-around performance for that particular airplane tips the risk balance in favor of having it available.

There is no point to use autofeather when there is no great benefit of that because:
Multiplying the phases where you rely on it, is multiplying the risk to forget it armed, and this feature being forgotten has been a source of incidents that could have turned into dramatic accidents.
Please refer to the following very interesting Website:
https://assets.publishing.service.gov.uk/media/5422ed34e5274a1317000179/dft_avsafety_pdf_500839.pdf
You may read:

The commander's decision to delay the selection of climb power was the initiating event in the sequence which started with the failure of the first officer to disarm the autofeather systemand allowed a train of events to occur which could have had more serious consequences. It is clear that had the standard practice of 'autofeather disarm - select climb power' occurred at the normal point, this incident would not have happened

Clearly the objective of limiting the use of this feature, is to reduce human factors consequences, while keeping its advantage in the most critical phase.

It could also be that Bombardier decided it wasn’t worth the cost of certification. An engine failure immediately after or during go around was unlikely, just like moments after V1. However, the certifying agencies do not give them any choice for the later.


text:
Is it legal to install a TACAN receiver on a civilian aircraft?
It is legal to use the DME from a VORTAC station (see Is it possible for a civilian to use a VORTAC station? for details), but is it also legal to use the TACAN for bearing determination?
This would probably require different equipment, which is not typically installed on a civilian aircraft (see Why doesn't civil navigation equipment support TACAN? for details). But the answers to that question do not mention whether or not it is legal to install and use this equipment.

You can buy and legally use TACAN equipment for GA aircraft from Bendix/King, but the system is expensive (close to US\$100,000), relatively heavy and requires rack-mounting.

They really went all-out in their sales copywriting: "An antenna is a part of this system, to maximilize the range of this system. Thanks to TACAN you will be able to control the pitch angle, and many other."

What's really funny is that the size and weight of the thing hasn't changed much since I worked on CH-46's in the Marine Corps.

Isn't € 1 373,96 in American format €1373.96, so we'll under $2000?


text:
How does fuel economy quantitatively depend on altitude for a jet?
We already have a few questions and answers about the qualitative effects of altitude on fuel economy:

How do I determine the best altitude to fly for fuel economy?
Why does fuel consumption decrease with increasing aircraft altitude?
What is the relation between an airplane's altitude and the drag it is experiencing?
Why do jet engines get better fuel efficiency at high altitudes? (contains formula for thermal efficiency)

I am interested in understanding how the changes in pressure and temperature at different altitudes affect the fuel economy of a turbofan powered aircraft quantitatively. In the end, I would like to make a plot of relative fuel economy vs. altitude that takes all of these effects into account, but I don't know how to quantitatively combine these effects.
Some notes:

By fuel economy I mean fuel required per distance traveled, not time.
I am not interested in absolute numbers for the fuel per distance, which would require specifying a particular aircraft. I am only interested in how the fuel per distance figure would relatively change with altitude, e.g. normalized to 1 at sea level.
I assume ISA (International Standard Atmosphere) profiles for pressure and altitude.
I assume the no wind case. Different winds at different altitudes will of course have an effect on the result, but it is easy to take this into account after the no wind case is understood.
Let's assume a typical climb profile for a short- to medium-haul jet airliner: 250/280/0.78

You can see that the TAS increases until reaching Mach 0.78, then decreases due to the lower temperatures causing a lower speed of sound and then remains constant above the tropopause. I am particularly interested in how the fuel economy will behave around these altitudes.

Fuel economy per hour, per mile? Set speed or set AoA?  A lot depends on gross weight. My flight planning cruise tables for a 737-400 indicate that best economy is found around 2000-5000 feet below service ceiling when operating m.74-m.78.

@MaxPower It's in the question: "By fuel economy I mean fuel required per distance traveled, not time." and "Let's assume a typical climb profile for a short- to medium-haul jet airliner: 250/280/0.78" (that is 250KIAS below FL100, 280KIAS above, until reaching Mach 0.78).

Sorry, I missed that. You still have many issues controlled by the specifics of each aircraft and even loading of each aircraft. You specify mach.78 but some jets have a critical mach below .78, and ground speed will change with temperature even at a set mach. Engines can be designed for any altitude. Similar for wing loading.

I am adding this community wiki answer to show my current state of research and to provide the plot vs. altitude, which I will update when I learn more. Comments are welcome.
Thermal Efficiency
From this answer by Peter Kämpf, we know that the thermal efficiency for jet engines is given by
$$ \eta = \frac{T_\text{max} - T_\text{amb}}{T_\text{max}} $$
where $ T_\text{amb} $ is just the ambient temperature (from ISA) and $ T_\text{max} $ is the temperature resulting from the combustion. If I understand the answer correctly, this should be about 1100 K above ambient temperature, so I am currently using this term to describe the impact of thermal efficiency on fuel economy:
$$ \epsilon_\text{T} \propto \frac{1}{\eta} = \frac{T_\text{max}}{T_\text{max} - T_\text{amb}} = \frac{T_\text{amb} + 1100 \, \mathrm{K}}{1100 \, \mathrm{K}} $$
I am not sure if the increase in temperature of 1100 K is constant with altitude, so please correct me if this is wrong.
Drag
From another answer by Peter Kämpf, we know that induced drag is proportional to dynamic pressure
$$ q = \frac{v^2}{2} \cdot \rho $$
with $ v $ being the TAS and $ \rho $ the density (known from ISA). Since work required to overcome the drag per distance is proportional to the force, the fuel economy should just scale with
$$ \epsilon_\text{drag} \propto \text{TAS}^2 \cdot \rho $$
Propulsive Efficiency
From this answer we know that the propulsive efficiency for a jet engine is given by
$$ \eta_p = \frac{2}{1 + v_e / v_0} $$
where $v_e$ is the exhaust velocity and $v_0$ is the TAS. As far as I could find, there is no straight-forward way to relate $v_e$ to altitude and temperature. For the moment, I added the propulsive efficiency for a high-bypass jet engine from the following plot from Wikipedia:

(image source: Wikipedia)
Summary
For the combined (relative) fuel economy term, I just multiply all previous terms:
$$ \epsilon = \epsilon_T \cdot \epsilon_\text{Drag} \cdot \epsilon_\text{Prop} $$
The following plot now shows the relative fuel required per distance. Each curve has been normalized to 1 at sea level.

The propulsive efficiency dominates as long as the TAS is increased. Afterwards, the lower drag dominates. The overall fuel use is almost half as low at high cruise altitudes compared to sea level.

Watch out. The “drag is proportional to dynamic pressure” is sloppy wording from Peter. It is a function of dynamic pressure, but there is a dynamic pressure where drag is minimal. And there is another indicated airspeed a little above that where the power of drag (drag times speed is the rate of energy loss due to drag) and these only depend on the dynamic pressure, which means indicated airspeed, and not altitude (much; see that Peter's answer). But the conclusion that drag is basically constant is correct.

Your conclusion is actually correct for propeller-driven aircraft. With decent constant speed propeller the propulsive efficiency, that is fuel consumed for unit of energy actually given to the aircraft is fairly flat. And because constant drag means constant energy for travelling given distance, the fuel consumption varies little with altitude. That is, however, not the case with turbojet and turbofan engines. Their propulsive efficiency increases quite significantly with forward speed. Unfortunately I didn't find good quantification anywhere.

@JanHudec Thanks for your comments. I knew I was missing at least one term. I will have a closer look at the jet propulsive efficiency...

the problem is that good data is difficult to find. Long ago I found some NASA lectures with a Java applet that is supposed to calculate this. The four links I wrote down are: https://www.grc.nasa.gov/WWW/k-12/airplane/specth.html, https://www.grc.nasa.gov/WWW/k-12/airplane/turbfan.html, https://www.grc.nasa.gov/WWW/k-12/airplane/ngnsim.html and https://www.grc.nasa.gov/WWW/K-12/airplane/EngineTheory.pdf. Back then I didn't see sources of the applet and I didn't get around to read through the paper, but sources seem to be there now.

@JanHudec I added propulsive efficiency now, but I'm not sure if the graph I found is useful for this comparison.

it would certainly be better to have some theory or at least empirical formula for it, but it is enough to show just how much jet engine efficiency increases with speed. Would be useful to accompany with plot/formula for the $\frac{\rho}{\rho_0}$ that gives the relationship between velocity and indicated airspeed (equivalent airspeed, actually) to see how quickly the velocity increases with altitude for given indicated airpspeed.

@JanHudec Something like that is plotted in the question: TAS and Mach as a function of altitude, where IAS is from the climb profile 250/280 and then later fixed Mach.

I can not see a comprehensive answer to the first question but will answer the next three

Why does fuel consumption decrease with increasing aircraft
altitude?

Because of an improved ratio of true speed to total drag. See question 2

What is the relation between an airplane's altitude and the drag it
is experiencing?

With reduced air density more speed is required to create the same dynamic pressure. Both lift and drag are direct functions of the dynamic pressure. The limit on altitude is a combination of Angle of attack and airspeed. Absolute ceiling is determined by gross weight, maximum lift coefficient(high AoA) and maximum dynamic pressure.(most often limited by critical mach and air density) Best altitude is found near the optimum angle of attack ceiling. Which results in just under critical mach and a dynamic pressure where lift=gross weight with an AoA of best lift to drag ratio. This will be below the absolute ceiling because best L/D AoA is generally much lower than the maximum lift AoA.

Why do jet engines get better fuel efficiency at high altitudes?
(contains formula for thermal efficiency)I will answer

Simply they are designed for best performance at high altitude and artificially throttled at low altitude. Because they are designed to compress cold low density air at high altitude they would overheat or have mechanical stress failures if operated at maximum near sea level. The biggest limit in gas-turbine design is the thermal limits of the material used in the high turbine first stage.(strength at a temperature) Operating below maximum is less efficient, this is a common thermodynamics/entropy engineering problem, the larger the difference in energy the more efficiently that energy can be harnessed(converted to another form) or transferred in or out of a system. This is also a limit in steam engine design, higher temperature boilers and higher pressure steam make them more efficient but material properties and safety are the limits with common heat sources.

Climb and decent roughly cancel out. The climb adds potential energy and the decent converts potential energy into kinetic energy. Any difference comes down to individual aircraft/engine design and operation parameters, and so is difficult to generalize.

The ratio of true speed to total drag is, in itself, irrelevant. Since the drag ends up being the same, the work for the same distance also remains the same (because work is force times distance). If the efficiency of the engine is independent of speed, the fuel consumption is mostly independent of altitude, which is actually true of piston-powered propeller planes! Turbojet and turbofan engines, and to some extent even turbofan engines are however more efficient at higher speed (the ram pressure increases their effective compression ratio), and that is the reason.

@MaxPower Thanks for the answer, but I think you misunderstood my actual question. The "questions" you answered are other related questions, which already have answers. I was interested in how to combine things quantitatively.


text:
Why is the landing gear lever not set to OFF on the ground?
Several older airliners (e.g. Airbus A300, Boeing 737 until the NG series, Boeing 747 until the -400 series) have landing gear levers with three positions: UP, OFF and DOWN. OFF (or NEUTRAL for Airbus) means the hydraulic pressure is removed from the system because the gear is held in place mechanically. See also Why does the landing gear control have 3 positions (up/down/off) on Boeing 737-800?
The gear lever is typically set to the OFF position after the gear is retracted during the takeoff and the system therefore remains unpressurized during the flight:

Pilot Monitoring After flap retraction is complete:
[...]

Set the landing gear lever to OFF after landing gear retraction is complete.

(Boeing 737 NG FCOMv1 NP.21.46 - Normal Procedures - Amplified Procedures)
The gear is then lowered by selecting DOWN before landing, but it is not selected back to OFF after the landing gear is extended. I can understand why this would not be done during the landing phase (redundancy by keeping the gear pressurized down, reduced crew workload during highly critical phase of flight), but why is the lever not selected back to to OFF after landing?
Every procedure I can find lists the gear lever as DOWN on the ground:

Landing Gear Lever:

Check the L/G lever DOWN and the EMRG retract pb protected.

(Airbus A300 FCOM - Cockpit Safety Inspection)

LANDING GEAR lever .....................................................................DN
Verify that the green landing gear indicator lights are illuminated.
Verify that the red landing gear indicator lights are extinguished.

(Boeing 737 NG FCOMv1 SP.6.1 - Supplementary Procedures - Electrical Power Up)
Hydraulic pressure is not required to keep the gear down:

Landing Gear Extension
When the LANDING GEAR lever is moved to DN, hydraulic system A pressure
is used to release the uplocks. The landing gear extends by hydraulic pressure,
gravity and air loads. Overcenter mechanical and hydraulic locks hold the gear at
full extension. The nose wheel doors remain open when the gear is down.

(Boeing 737 NG FCOMv2 14.20.2 - Landing Gear - System Description, emphasis mine)
And furthermore, hydraulic pressure is lost anyway once the engines (and electrically powered hydraulic pumps) are switched off on the ground. So why would you want to keep the gear pressurized down on the ground until shutdown and again during startup?

Because it has to move through UP to get there

@ratchetfreak Why? The OFF or NEUTRAL position is in the center. At least on the A300, it is possible to move from DOWN to NEUTRAL: "On ground, L/G lever can be moved from DOWN to Neutral position and vice versa, but interlocks prevent an inadvertent UP selection." (from here). Is there something in the Boeings that would prevent selecting OFF from DOWN?

@ratchetfreak No, after takeoff you can stop at Off on your way to Up -- the detent is there. It's pointless, but possible.

The NG 737s have the Off position; the Max's don't.

Because there's no advantage to doing so (since the pumps will soon be off anyway, as the question notes), and this would create the risk that the pilot may overshoot the Off position to the Up position, which could, on a bad day, lead to an inadvertent gear retraction and major damage. The over-center locks are good, but do you really want to bet big bucks that the combination of 3,000 psi + a bad bump in the pavement will never produce an illustration of Murphy's Law?
Okay, there is one other safety to prevent overshooting, but that solenoid can fail, and if it does, watch out! There is simply zero need for this.
Also, depending on the plumbing, you might remove hydraulic pressure from the nosewheel steering by doing this (though I'd have to check the diagram to be sure).

OK, I may have over-estimated the advantages of disconnecting the hydraulics. "you might remove hydraulic pressure from the nosewheel steering" -> that would be really good reason not do disconnect on ground. My hydraulics diagram in the FCOM indeed connects nose wheel steering via the landing gear, but it does not say if it is actually shut off when the lever is in OFF.

Hydraulics to the landing gear are shut off in flight for a couple reasons.

Reduction of internal leakage of the hydraulic system. Internal leakage increases pump power extraction from the engines very slightly increasing fuel burn. Also internal leakage consumes available hydraulic power that can not be used for other tasks. Available hydraulic power is already reduced at cruise and descent due to lower engine speeds (max flow from engine pumps is linearly proportional to engine speed).
Fault isolation. If a component starts to leak in flight, or there is a tire burst that damages hydraulic lines, it can't drain the entire hydraulic system over a multiple hour flight.

These concerns just aren't very meaningful on the ground. A small leak is unlikely to deplete you hydraulic system during the limited ground time, and a large leak you want to know about while you are still on the ground. Fuel burn over 20 minutes of ramp time from internal leakage is going to be completely negligible, and, outside of speed brakes, there are no other large consumers of hydraulic power on the ground.
While you taxi, bumps can place an unlocking force onto the gear (this is why some aircraft have bungees to add locking force), so having a hydraulics on as a back up isn't a bad thing either, just in case. While parked at the gate all loads are basically straight down, no drag or lateral loads impacting the gear.
The answer by @Ralph J is also true. Why run the risk of accidentally raising the gear on the ground? Why add to the pilot workload during a high workload time? Brakes and steering are also generally disabled when the handle is in the "off" position as well.
That being said, some electronically controlled landing gear do switch their extension/retraction actuators to neutral once down and locked automatically.


text:
How is fuel mass measured in airliners?
Large airliners display the amount of fuel in each tank as a mass (usually in kg or lbs) and not as a volume. This makes sense since mass is important for weight&balance considerations and the fuel mass is proportional to the amount of available energy.
Based on a discussion in the comments under this answer, I am wondering: how is the fuel mass measured?

On the Airbus A320 family, the fuel volume is measured via capacitance probes and then converted to mass by also measuring the fuel density:

The FQI [Fuel Quantity Indication] system calculates the fuel quantity based on values taken from probes in the tanks. The probes measure the level of the fuel in the tank, as a consequence of changing capacitance due to the amount the probe is immersed. This allows the determination of the fuel volume in the tank.
Yet the information that is needed by pilots is the quantity of fuel on board expressed as a weight. The translation of fuel volume into fuel weight is performed by the FQIC using the fuel density measured by specific devices in each wing tank [...]

(Airbus Safety First - Fuel monitoring on A320 Family aircraft)
The FCOM lists the sensors involved in this measurement:

The FQI system comprises : [...]

A set of capacitance probes in each tank to measure fuel level and temperature.
one densitometer (cadensicon) sensor in each wing inner tank permitting the calculation of the fuel quantity.
one Capacitance Index Compensator (CIC) in each inner tank giving the dielectric constant of the fuel in case of cadensicon failure.

(Airbus A320 FCOM - Fuel Description - Fuel Quantity Indication and Level Sensing)
Using the densitometer (or the backup CIC), the conversion from volume to mass is performed by the FQIC. The ECAM can then display fuel temperature and mass as calculated by the FQIC.

Additionally, there is an independent system sensing low fuel level, which would alert the pilots even if something goes wrong with the normal FQIS:

In addition to the sensors and probes feeding the FQI system, each wing tank is equipped with three independent dedicated low level sensors. These sensors are located in such a way that they become dry when the remaining fuel in the tank is approximately 750 kg. If two sensors in the same tank remain dry for more than 30 seconds, a low level alert triggers in the cockpit (fig.2).
The low level sensors are fully independent from the Fuel Quantity Indication, and are different in that they:

Do not provide pilots with a continuous indication of the fuel quantity in the wing tanks, but only the signal that the fuel level has reached below 750 kg (threshold crossed).
The information provided to pilots in the form of the low level alert results from a physical measure (sensors dry or wet) rather than from a calculation.

(Airbus Safety First - Fuel monitoring on A320 Family aircraft)

Pretty much all transport airplanes use capacitance systems.

@JohnK Yep, that's what I read, too. The 777 seems to be an exception using ultrasonic sensors. But here I'm more interested in how they convert the volume to mass. That seems to be different between aircraft types.

So in short, fuel mass is not measured, it is computed from volume and other factors.  Just as you could compute the mass of fuel in your car's tank by knowing how many gallons/liters are in it, and the specific gravity of gasoline or diesel.

@jamesqf Exactly, except that the specific gravity is also measured in this case since it strongly depends on temperature (and of course also the type of fuel, like Jet A vs. Jet A-1).

@Bianfable: Sure, just as the specific gravity of gasoline varies, too, depending on temperature, specific blends from the refinery, additives such as ethanol, &c.  So it's much easier just to measure gallons, and a really dumb gauge that just shows fuel level.

fyi I've just asked How does a fuel density sensor work? (densitometer, cadensicon) and quoted your answer.


text:
What is this component below the fuselage of the Fokker F28 used for?
There is a long straight component with three attachment points below the fuselage of the Fokker F28 Fellowship:

(image modified from Wikimedia)
It looks like it is installed on both sides of the fuselage:

(image modified from Wikimedia)
What is this component used for?

Possibly an ADF sense antenna?

those are antenna supports, probably for a radio navigation system, that hold the thin antenna rod in a straight line and keep it from vibrating in flight.


text:
What is "spatted undercarriage"?
The Socata TB-family wikipage lists the TB-9 and TB-10 as having "spatted undercarriage". What does that mean? From the pictures it appears to be a normal tricycle-gear-setup so I assume it's not referring to the layout of the landing gear.

If I'm not mistaken, a spat is a fairing for the gear to make it more aerodynamic.

An aircraft wheel fairing, commonly called a wheel pant or spat or, by
  some manufacturers, a speed fairing. Wikipedia

Thanks! I visited that page, but somehow I missed that. I always thought they were simply called "wheel fairings", you learn something new everyday :)

@roe it doesn't seem like a very common term, more informal and less used in literature. might be limited to the united states as well.

"Spatted undercarriage" refers to streamlined fairings on fixed landing gear. They're also known as "wheel pants", although when it comes to comparisons with clothing "spats" is certainly more accurate, as they only cover the top of the wheel, not the entire gear leg.
The term is semi-common in the US, but I've mostly heard (and used) "wheel pants" when talking about landing gear fairings.
Spats / wheel pants / fairings on an AA-1's landing gear:

Spats on shoes on Charles Sumner:

license: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/3.0/


text:
If an airplane touches down short of the runway, and doesn't reach it, is it still a runway excursion?
In light of the two recent incidents where aircraft landed roughly 1000 ft short of their intended runways, Air Canada AC-624 at Halifax and Asiana OZ-162 at Hiroshima, I started thinking the definitions used when labeling them.
At least in the first case, the airline labeled it a runway excursion. Given that the airplane never actually exited the runway (it touched down short and slid onto the runway), I'm a little confused, what is the definition of a runway excursion? If an aircraft never actually comes in contact with the runway, is it still a runway excursion? At what point does it simply become an off airport landing?

In the case of the OZ flight, it technically would be CFIT followed by a runway excursion, I suppose. Same for their crash at SFO. In both cases, they crashed into the ground short of the runway (CFIT,) slid onto the runway, then skidded off of the runway (runway excursion.)

I still think "runway excursion" sounds like a nice Sunday afternoon stroll.

It depends.
Skybrary says

The Runway Excursion category includes also two types of occurrences which do not fit the ICAO ADREP definition for a runway excursion, however considered appropriate for inclusion due to the commonality of a number of causal and contributory factors and/or mitigation approaches:

An aircraft attempting a landing touches down in the undershoot area of the designated landing runway within the aerodrome perimeter.
A runway or taxiway other than the designated one is used for a take off or a landing.

My emphasis.

The European Commercial Aviation Safety Team (ECAST) in the European Strategic Safety Initiative (ESSI) use the Skybrary definition when writing about Runway Excursion Prevention

Articles within the Runway Excursion Category in SKYbrary:

Overrun on Take Off
Overrun on Landing
Directional Control
Undershoot on Landing
Wrong Runway

My emphasis.

In an article "Language and safety issues" in Eurocontrol's magazine in 2011, Bert Ruitenberg wrote:

The whole idea of putting a label such as “runway incursion” or
  “runway excursion” on a safety occurrence is to make it easier to
  file the data from the event somewhere and to compare it with
  similar occurrences. With the definitions above, a take off from a
  taxiway would be considered as a Runway Excursion when the
  Skybrary definition is used, but not with the other RE definitions
  (and rightly so, I say, better label it a “taxiway take off”). Moreover,
  in the ICAO definition, the B777 undershoot at Heathrow
  would not be a RE, but with the Skybrary definition it would be.
  Dear Safety Managers of the world, there still is a lot of work to
  be done...

all_licenses: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/3.0/